Three-axis attitude determination systems for current space and launch vehicles use separate sensor suites involving different combinations of expensive units (earth sensor, sun sensor, gyro, star tracker, etc.) for spinning or non-spinning applications. For example, the spin axis attitude determination of a spinning space vehicle has been traditionally accomplished by using a combination of sun and earth horizon sensors. A sun sensor measures the sun illumination spike and the elevation angle with respect to the spin axis of the space vehicle. Sun position and elevation information establishes a geometric sun position cone with a half cone angle equal to the elevation angle on which the spin axis lies. A similar earth elevation cone can be constructed based on the earth sensor measurements. The intersection of the sun position cone and earth elevation cone provide the inertial attitude information for determining the spin axis of the rotating vehicle. The cost of such a system can be appreciable. It is desirable to find cost effective alternatives to the use of expensive sun and horizon sensors.
Some researchers have pursued three-axis attitude determination, using GPS (Global Positioning System), based on interferrometry of a spatially separated multi-patch antenna system for inertially stationary platforms. A major disadvantage of this interferrometric approach is the requirement of large antenna baselines. This requirement makes the product bulky and heavy. Moreover, it cannot be used on a spinning system. Accordingly, it would also be desirable to find alternative products that are small, light, and power efficient. Furthermore, it would be desirable to provide a single attitude and navigation sensor for both spinning and non-spinning systems, e.g., for land, marine, airborne, and space applications. Potentially, such a unified GPS-based sensor would provide significant size, weight, power, and/or cost advantages.